This invention relates generally to turbine engines, and, more particularly, to slot cooled ring combustors for turbine engines.
A turbine engine includes a compressor for compressing air which is mixed with a fuel and channeled to a combustor wherein the mixture is ignited for generating hot combustion gases. The combustion gases are channeled to a turbine, which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight or to power a load, such as an electrical generator. Increased efficiency in gas turbine engines is accomplished at least in part by an increase in the operating temperature of the combustor. A limitation on the operating combustor temperature is a temperature limitation of combustor liner material.
Thin film convection cooling can be used to cool a combustor liner. With such cooling, a protective film boundary of cool air flows along an inner surface of the liner. The cool air flowing along the combustor liner inner surface forms a is protective boundary between the liner and the hot gases, and insulates the liner from hot combustion gases. See, for example, U.S. Pat. No. 4,259,842. Even with such cooling, however, the liner materials absorb heat. Over time, thermal creep and low cycle fatigue increase in the liner.
A thermal barrier coating also can be applied to inner surfaces of the combustor liner for providing thermal insulation against combustion gases. Thermal barrier coatings reduce an amount of cooling air required for a given combustion gas temperature, or allow an increase in a combustion gas temperature for increasing efficiency of the engine. See, for example, U.S. Pat. No. 5,960,632. Typically the thermal barrier coating is applied uniformly across the combustor liner with a thickness of 0.01 inches or less. Such a uniform thickness prevents the thermal barrier coating from undesirably building-up to potentially obstruct the flow of cooling air. However, the combustor liner materials still absorb heat, and thus, combustor assemblies are still subjected to thermal strains including creep and low cycle fatigue.
In an exemplary embodiment, a combustor includes a combustor liner s with a thermal barrier material that has a thickness selected to minimize heat absorption. In the exemplary embodiment, the combustor includes a combustion zone formed by annular outer and inner supporting members and respective inner and outer liners. The inner and outer liners each include a series of panels and a plurality of cooling slots. The panels are arranged in steps relative to one another and form a stepped combustor liner surface. The plurality of cooling slots are formed by overhanging portions of the inner and outer liner panels. At least one portion of the combustor liner has a thermal barrier material with a thickness greater than 0.01 inches. In the exemplary embodiment, at least the outer and inner liner panels adjacent an inlet of the combustor have a thermal barrier material with a thickness greater than 0.01 inches.
As a result of the additional thickness of thermal barrier material applied to at least a portion of the combustor liner, the combustor liner material absorbs less heat, and therefore, at present day operating temperatures, the combustor may be operated at higher temperatures. Because the operating temperature is reduced, low cycle fatigue within the combustor is also reduced which, in turn, extends an operating life cycle of the combustor assembly.